1. Technical Field
The present invention relates generally to gas turbine engine turbine shrouds and, more specifically, mounting and sealing of such shrouds.
2. Background Information
A conventional gas turbine engine typically includes a compressor, combustor and turbine, both rotating turbine components such as blades, disks and retainers, and stationary turbine components such as vanes, shrouds and frames routinely require cooling due to heating thereof by hot combustion gases.
The high pressure turbine (HPT) stages typically maintain a very small tip clearance between turbine blade tips and shrouds surrounding the tips. Shroud supports maintain the shrouds in desired position relative to the rotating blade tips to control clearances between the shrouds and blades. The tip clearance should be made as small as possible for good efficiency, however, the tip clearance is typically sized larger than desirable for good efficiency because the blades and turbine shroud expand and contract at different rates during the various operating modes of the engine.
The shroud is conventionally an assembly of a plurality of gas turbine engine stationary shroud segments assembled circumferentially about an axial flow engine axis and located radially outwardly about rotating blading members and defines a part of the radial outer flowpath boundary around the blades.
The shroud segment and shroud assembly must be capable of meeting the design life requirements selected for use in a designed engine operating temperature and pressure environment. To enable current materials to operate effectively as a shroud in the strenuous temperature and pressure conditions as exist in the turbine section flowpath of modem gas turbine engines, it has been a practice to provide cooling air to a radially outer portion of the shroud. However, as is well known in the art such cooling air is supplied at the expense of engine efficiency. Therefore, it is desired to conserve use of cooling air by minimizing leakage into the flowpath of the engine of cooling air not designed in the engine.
Composite and, in particular, ceramic matrix composite (CMC) materials have been suggested for use in shroud segments because they have a higher temperature capability than the metallic type materials currently in use. However, such materials, forms of which are referred to commercially as a ceramic matrix composite (CMC), have mechanical properties that must be considered during design and Application of an article such as a shroud segment. CMC type materials have relatively low tensile ductility or low strain to failure when compared with metallic materials. Also, CMC type materials have a coefficient of thermal expansion (CTE) significantly different from metal alloys used as restraining supports or hangers for shrouds of CMC type materials. Therefore, if a CMC type of shroud segment is restrained and cooled on one surface during operation, forces can be developed in CMC type segment sufficient to cause failure of the segment.
Generally, commercially available CMC materials include a ceramic type fiber, for example SiC, forms of which are coated with a compliant material such as BN. The fibers are carried in a ceramic type matrix, one form of which is SiC.
The turbine shroud directly affects overall efficiency or performance of the gas turbine engine due to the size of the tip clearance. The turbine shroud additionally affects performance of the engine since any compressor discharge and/or bleed air used for cooling the turbine shroud is therefore not used during the combustion process or the work expansion process by the turbine blades and is unavailable for producing useful work.
Accordingly, it is desirable to control or reduce the amount of bleed air used in cooling the turbine shroud for maximizing the overall efficiency of the engine. It is also desirable to use CMC materials in the shroud because they have a higher temperature capability than the metallic type materials currently in use.